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16.89J / ESD.352J Space Systems EngineeringSpring 2007

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MINERVACritical Design Review

16.89May 8, 2000Department of Aeronautics and AstronauticsMassachusetts Institute of Technology

Mission Statement.

Establish an enabling space infrastructure that will support the exploration of Mars.

2

Agenda

IntroductionGeneral mission overviewDetailed designSystem level issuesLessons learned and conclusions

Introduction

4

Motivation for MissionDramatically enhance the value of future Mars missionsInfrastructure at Mars provides major increase in science return

Pathfinder: 30 MB/sol MINERVA: 10 GB/solSupport for up to 10 Mars Surface Elements (MSEs)Accurate location information

Robotic mission designers can focus on science missionEnhanced probability of mission successMore science for the taxpayer’s dollar!

User Needs

MINERVA system shall provide enabling infrastructure to support exploration of Mars.The infrastructure shall provide Mars Surface Elements (MSEs) with:

Communication services between Mars surface and Earth Ground Stations (EGS)Their position on the surface of Mars, without imposing additional design constraints on MSEs.

Requirements Flow Down

MINERVA (M)

Earth Based (E)Mars Orbiting (S)

Payload (P) Bus (B)

Program (Z)

General Mission Overview

8

Design Summary

Mars-orbiting constellationNumber of spacecraft: 4Number of orbit planes: 2Altitude: 2000 kmInclination: 27°

Spacecraft wet mass: 470 kgSystem cost: $297.9 M

Drivers: software development, launch

9

Launch

4039 dia(159.0)

3750 dia(147.6)

775(30.5)

4366(171.9)

8893(350.1)

912 dia(35.9)

15oLaunch date 18 Aug 2007Launch window ± 1 sec, every

1 sidereal day from3–18 Aug 2007

Launch site Cape CanaveralAir Station

Launch vehicle Delta IIIVehicle provider BoeingTotal mass 1974 kgShared payload Possible, but not

necessaryConfiguration Four stacked

spacecraft

Image by MIT OpenCourseWare.

10

Transit Overview

Earth (start) Mars (start)

Earth (final)

Launch 18 Aug 2007Departure burn T+ 0d 3:23Separation T+ 0d 3:29Deploy arrays T+ 0d 6:01Initial checkout T+ 0d 6:05Alignment burn T+ 2d 16:39Correction burn T+ 122d 16:00Insertion burn T+ 285d 14:29Circularization T+ 290d 8:22Deploy antenna T+ 290d 8:24Test/calibration T+ 296d 12:00IOC 10 Jul 2008

Mars (final)

Separation &Deployment

Spin-up &Insertion

CorrectionBurn

AlignmentBurn

11

Day in the Life: Positioning

.

DSN

Two-way Doppler trackingover 10-hour DSN pass

180 measurements per day:• Two-way ranging• Two-way Doppler tracking

Coarse estimate: • 10s km immediate• Best estimate >1 km • Best obtained in 3 hr• Update period 35 min

Daily post-processing: • 100 m accuracy• 35 min update rate

12

Day in the Life: Communication

.DTE - 1

DTE - 2DSN

13

Day in the Life: Communication

.DTE - 1

DTE - 2DSN

14

End of Life: DisposalSatellite has capability to insert into a disposal orbit

Boost to 2150 km altitudeRequires only 40 m/s ΔV

Allows constellation replenishment

Detailed Design

16

Design Iteration Process

Integrated Concurrent Engineering (ICE)• # S/C

• Altitude

• Inclination

• # Orbit planes

• Earth parking ...orbit

Design Vector

OrbitsPayload

Bus

Systems• System cost

• Cost per function

Launch

• Max cone angle

• S/C mass

• Exhaust ...velocity

• Availability

• Revisit time

• Max eclipse ...time

• Total ΔV

• P/L mass

• Cost

• Power

• Lifetime ...performance

• Total mass• 1st unit cost

• Launch cost

• RDT&E cost

17

ICE Design Sessions

Identified best launch scenarioDirect to Mars transfer over LEO parking orbitSwitch to chemical propulsion over electric

Identified best constellation altitude2000 km for four spacecraftMinimizes system cost

Discovered minimal cost saving with three spacecraft

Sacrificing availability and robustnessTweaked inclination orbit

Significantly reduces maximum revisit time

Detailed Design:Orbit Analysis

19

Orbits Requirements

M004 MINERVA shall have a maximum revisit time of less than 3 hours.

M005 MINERVA shall provide a coverage of ± 15° latitude band around the equator.

S001 Constellation shall have a minimum of 2 spacecraft in view of the Earth at all times.

S007 MINERVA shall have a crosslink availability of 90%.

20

Earth Interplanetary MarsΔV

(km/s)Time(d h)

ΔV(km/s)

Time(d h)

ΔV(km/s)

Time(d h)

Chemical 3.80 2d 17h 0.17 282d 23h 1.60 3d 17h

Electric 7.38 421d 14h 5.66 323d 3h 2.63 150d 1h Using 185km parking orbit

Transit Method Trade StudyProposed methods for the interplanetary segment

Chemical propulsionElectric propulsion

Design discriminators from an orbit standpointTotal ΔV for all phases of the missionTime of flight for transit to Mars

21

Transit Method Trade StudyEarth Interplanetary Mars

ChemicalPropulsion

ElectricPropulsion

22

Transit Method SelectionConsiderations

Chemical propulsion provides fast transfer for smaller ΔVElectric propulsion is more benign

More time to react to problemsSmaller forces exerted during maneuvers

Conclusion: from orbit standpoint, chemical propulsion is recommendedOther groups are involved in this trade

Bus GroupSystem Group (Cost)

23

Launch OpportunitiesEach Earth-Mars launch window has a slightly different ΔV requirementThe MINERVA design can accommodate all three launch opportunities investigatedThe launch window in 2009 may be used as a backup opportunity, with system IOC on 23 Sep 2010

Launch Departure ΔV Capture ΔV Time of Flight

2005 3.726 km/s 1.742 km/s 278d 15h 35m

2007 3.799 km/s 1.601 km/s 290d 8h 22m

2009 3.712 km/s 1.753 km/s 278d 21h 54m

24

Delta III Launch Sequence

1.

2.

3.

4.

5.

T+ 0:00 Launch 1T+ 1:19 Solid drop (6) 2T+ 2:37 Solid drop (3) 3T+ 3:44 Jettison fairing 4T+ 4:29 Stage 1 separation 5

T+ 4:41Stage 2 burn, i=28°ΔV = 4.628 km/sDuration = 8.27 min

6

T+ 16:00 Collision avoidance run 7

T+ 28:17Stage 2 burn, i=23.45°ΔV = 0.700 km/sDuration = 35 sec

8

6.

7. 8.

25

Transit - Departure

T+ 3:23:20

Departure burn (second stage) ΔV = 3.799 km/s Duration = 5.59 min

T+ 3:29:30 Start release sequence Interval = 50.15 min

T+ 6:01:00 Despin maneuver T+ 6:01:50 Deploy solar arrays T+ 6:05:00 Initial checkout T+ 2d 16:39 Depart Earth SOI

Fairing Jettison

Satellite Separation

Solar Array DeploymentSolar Arrays gimbaled about North-South axis

Cross-Link DeploymentDeploys on hinged boom

27

Transit - Rendezvous

T+ 2d 16:39

Alignment burn (four ACS thrusters) ΔV = ~0.020 km/s Duration = 48.2 sec

T+ 2d 16:45 Functional testing

T+ 122d 16:00

Correction burn (four ACS thrusters) ΔV = ~0.005 km/s Duration = 12.0 sec

T+ 285d 00:00 Upload precise position T+ 285d 01:00 Spin-up maneuver

T+ 285d 14:29 Arrive Mars SOI (29 May 2008)

28

Capture and Deployment

T+ 285d 14:29

Injection burn (main kick motor) ΔV = 0.167 km/s Duration = 2.1 sec

T+ 290d 08:22

Circularization burn (main kick motor) ΔV = 1.602 km/s Duration = 19.1 sec

T+ 290d 08:23 Despin maneuver T+ 290d 08:24 Deploy large antenna T+ 290d 10:54 All satellites in place

T+ 291d 12:00 Correction maneuvers (as necessary)

T+ 296d 12:00 Test and calibration T+ 326d 01:40 IOC: 9 July 2008

Earth-Antenna Deployment Full pointing capabilities using 2 DOF boom

Nominal mission configuration

Fully Deployed Satellite

30

Lifetime VisibilityEarth-Mars distance is periodic over 2.2 yearsExclusion zone of 19 days caused by line-of-sight intersection with the sun and its corona

31

Constellation Constraints

Recap of requirementsProvide coverage to a ± 15° latitude bandMinimum MSE to satellite availability of 50%Maximum revisit time of 3 hrs

Architecture constraintsAllow for line of sight communications between satellitesMinimum inclination of ≈ 30°

32

Trade Spaces

Coverage requirementsAltitudeNumber of satellitesInclination (restricted by the position determination requirement)

Constrained by the cross-link requirementsAltitudeNumber of satellites

Cost (looked at in ICE sessions)AltitudeNumber of satellites

33

Coverage Trade Space

Constrained by:Revisit time < 3 hrs50% availability

Variables:Number of satellitesInclinationAltitude

34

Cross-link Trade Space

Minimum altitude required for cross-linksSignal beams pass at least 200 km above the surface of Mars

Final Constellation Design

Walker-Delta patternCircular orbits2 Planes4 Satellites27° Inclination

36

Percentage of Time in View

Constellation provides >70% coverage in the ± 150 latitude bandReduced coverage

up to ± 650

37

Revisit Time

The maximum time between satellite passes is <30 minThe average time is <20 min

38

Contact Duration

On average, a satellite will remain in view for 50 minutes.

Final Constellation Design

Walker-Delta pattern4 satellites in 2 planesInclination of 270

Provides (± 150 Lat)Avg. revisit time < 20 minMax. revisit time < 30 minContact duration ≈ 50 min Availability > 70%3 satellites in view of EarthReduced coverage up to(± 600 Lat)

Single Satellite Failure

In the event of a single satellite failure, the constellation will be able to provide communication and navigation at a diminished level

Provides (± 150 Lat)Avg. revisit time < 45 minMax. revisit time < 100 minContact duration ≈ 50 minAvailability > 50%At least 2 satellites in view of Earth

Detailed Design:Payload Analysis

42

Payload Requirements

M001 MINERVA shall provide communication capability between MSEs and EGS for at least 10 continuous hours per day.

M002 MINERVA shall provide MSE position accuracy of 100 m (horizontal resolution) or less.

M003 MINERVA shall return MSE position determination daily with an update every 3 hours.

S005 Constellation shall return a minimum of 10 Gb/sol data rate to EGS.

43

Payload Requirements (cont.)

E002 EGS shall be able to resolve spacecraft orbit to an accuracy of 20 m in radial, along-track, and cross-track directions.

E003 EGS shall be able to upload spacecraft orbital element data and clock offsets at least once per day.

E008 Uplink from EGS to MINERVA shall have a BER of no greater than 10-9.

E009 Uplink from EGS to MINERVA shall have a data rate of at least 500 bps.

44


P001 Payload mass shall not exceed 50 kg.

P002 Payload shall use UHF for communication withMSEs.

P003 Uplink from MSE to MINERVA shall have a BER of no greater than 10-6.

P004 Payload shall have a downlink BER no greater than 10-6.

P005 Each satellite shall have a downlink data rate of at least 150 kbps from MINERVA to EGS.

45


P006 Payload shall dynamically allocate downlink data rate and uplink from MSE to constellation data rate.

P007 Payload shall provide 30 Gb storage for communication data.

P008 Payload subsystem shall use an on-board orbital propagator with an accuracy of 10 km for backup.

Payload Analysis:Communication

47

Communications Requirements

Communication systemRelay between Mars Surface Elements (MSEs) in the ±15° latitude band and the Earth. Exceed 10 Gb/sol of total data return

48

Communication System Overview

.

DSN

Cross-link

Three types of links

Earth-MINERVA Link

MINERVA-MSE Link

49

Antenna Types Analysis

Parabolic antennaOptimized for high gain (>20 dB) and low beamwidth (order of 15 deg or less)Has a lot of experience in space

Helix antennaOptimized for frequencies below 2 GHzBest suited for low gain and high beamwidthLight mass

Image removed due to copyright restrictions.


50

Antenna Types AnalysisPhased array antennaGenerates one or more beamssimultaneouslyChanges direction of the beam rapidlySweeps good gain over a largebeamwidth (e.g. 14 over 120°)No moving mechanical parts

Horn antennaOptimized for frequencies of 4 GHz or higherBest suited for low gain and high beamwidthHigh weight



51

1) Omnidirectional antenna

• Inefficient use of available power

Top Level Trade Analysis for the Communication System

.

DSN

Case 1: Integrating all links together in one antenna

52

2) Directional antenna

• Impossible to communicate between Mars and Earth at the same time ...(parabolic reflector and phased array antenna)


.

DSN


53

Conclusion:

• Integrating all links together is not the optimal solution


.

DSN


54

1) Using a helix type antenna or a parabolic antenna

• Not enough gain for that large beamwidth


.

DSN

Case 2: Integrating cross-link and MINERVA-MSE

55

2) Using a phased array antenna

• UHF phased array antenna have not been used for space ...communication


.

DSN


56

Conclusion:

• Integrating cross-link and MINERVA-MSE is not the optimal solution ...for this application


.

DSN


57


.

DSN

Case 3: Separating each type of link

One different type of antenna per link

58

Conclusion:

• Separating each type of link is the solution chosen


.

DSN

Case 3: Separating each type of link

59

Earth Ground Station Interface

Deep Space Network: 70-m vs. 34-m antennas

34-m: availability of Ka-band allows reduced satellite antenna size34-m: processing facilities located on the ground

Better thermal control - reduced system noiseSmaller operation cost

60

Modulation used

! BPSK R-1/2 Viterbi software decoding

" Standard deep space telemetry modulation format

Respects Shannon Limit

Figure by MIT OpenCourseWare.

elc

Stamp

61

Frequencies usedKa-band (32 GHz) for Earth-MINERVA link

Reduces the size of the antenna while keeping a high gainWill be supported by DSNAlso used during Earth-Mars transit

X-band (7 GHz) for cross-linkProvides good beamwidth without significantly influencing the antenna diameter (medium gain)Widely used in deep space missions

UHF (0.4 GHz) for MINERVA-MSE linkGood performance for omnidirectional antennas on Mars surfaceReduces necessary antenna mass on board MSE

62

Antenna Types Trade Analysis

MINERVA - Earth link: Parabolic antennaMars Earth distance: 50 - 400 million km

⇒ high gain requiredMINERVA - Mars link: Helix antenna

UHF 0.4 GHz to support existing assetsHigh beamwidth to improve coverage(77 deg at 2000 km altitude)

MINERVA cross-links: Parabolic antennaNecessity to use antenna for Earth link during Mars approach and as a backup

63

Payload HardwareAntennas and Transponders

MINERVA-Earth Link: Ka-(X)-band2.05 m parabolic, 130 W, 26.6 kg

MINERVA-Mars Link: UHFø 25 cm x 31 cm helix, 21 W, 2.9 kg

MINERVA Cross-Link: X-(Ka)-band2 x 50 cm parabolic, 5 W, 5.6 kg2 Omni-directional, 5 W, 0.3 kg

Total Mass: 35.4 kg

Other Hardware

Total Mass: 7.2 kg

Computer:RAD 6000, 5 kg

Used on Mars Pathfinder, Globalstar, ISS

Navigation Equipment:Ultra Stable Oscillator, 0.2 kgOther equipment:Switches, etc. 2 kg

Total Payload Mass: 42.5 kg

64

Cross-links5.9 kg

Earth link20 kg

Amplifiers6.6 kg

Mars link2.9 kg

Other2.2 kg

Computers5 kg

E a r t h l i n k

C r o s s - l i n k s

M a r s l i n k

3 S o l i d S t a t e A m p l i f i e r s

C o m p u t e r s

O t h e r

Payload Mass Breakdown

Total Mass: 42.5 kg

CommunicationsF.O.V. Verification

Use model to verify clear “lines of sight”between satellites, Mars and Earth

Payload Analysis:Position Determination

67

Position Determination Requirements

Position determination systemGather information to determine position of Mars Surface Elements (MSEs) in the ±15° latitude band With an accuracy of 100 mWith an average update period of less than 3 hours

68

ActivePassive

Positioning Design TradesPosition determination problem Active

Computational load

Infrared not proven

Radar Infrared

Single coverage for 2-DDouble coverage for 3-D

No requirement on user

Pros

Cons

One-way (GPS-like)

Unlimited #users

DopplerRange

Triple coverage for 2-DQuadruple coverage for 3-D

Time offset Frequency offset

Proven methods

Pros

Cons

Two-way

Quick 2-D positioningwith single satellite

Limited #users

Double coverage for 2-DTriple coverage for 3-D

Range Doppler

Transponders on user

Pros

Cons

ActiveActivePassive

One-way (GPS-like)

Range Doppler

Two-way

69

Position Determination Method

Vs

MINERVA satellite

MSE at(Φ,θ)

Knowntopology

R

Sphere at R from sat.

Roundtrip delay:

delay processing 2 tcRT Δ+=Δ

Sphere/sphere intersection

α

Cone at α from sat. velocity

Roundtrip Doppler shift:

λα )cos(2 SVf =Δ

Sphere/cone intersection

?

70

Ambiguity Resolution

Mars equator i

Satellite A at t1

L ~ 70 km / 5 min

Mars rotation:L= RM ωΜ ΔT cos(latitude)

Satellite A at t2 = t1 + k ΔT

(ΔT = 5 min)


(ΔT = 5 min)

Ambiguity resolution:Δx ~ 2Lsin(i)

Δx ~ 63 km / 5 min


(ΔT = 5 min)


Δx ~ 63 km / 5 min

Mars equator i

Satellite A at t1

L ~ 70 km / 5 min

Mars rotation:L= RM ωΜ ΔT cos(latitude)


(ΔT = 5 min)


(ΔT = 5 min)


Δx ~ 63 km / 5 min


(ΔT = 5 min)


Δx ~ 63 km / 5 min

71

Sources of Error

4 – 90 km errorPer km/hr

< 1 cm/sAssumed very slowMeasured with IMUs

MSE velocity

Absolute upper bound on accuracy

100 m – 10 km

20 m

Quick positioning: orbit predictionPost-processing:

orbit determination

MINERVAorbits

Corrected with time

~ 200 mMars topographyMSE altitude

Not limiting factor

< 1 cm/s< 1 mm/s

Integration time Sat. oscillator stability

Doppler error

Not limitingfactor

10 mCode chip rateRanging error

EffectMagnitudeProperties

Inte

rnal

Exte

rnal

72

Time to Get 100 m Accuracy Probability to reach 100 m accuracy (1 σ) within certain time:

0° latitude

73

Time to Get 100 m Accuracy Probability to reach 100 m accuracy (1 σ) within certain time:

15° latitude

Detailed Design:Software Analysis

75

MINERVA Software Components

Flight softwareTest, integration, and simulation software

Used to verify initial and updated flight software and during anomaly recoveryCost modeled in CERs

Operations softwareMission & activity planningMission controlNavigation & orbit controlSpacecraft operationsData delivery, processing, and archiving

76

Flight Software

”Estimation by similarity" technique used to estimate:

Source lines of code (SLOC)Software throughput requirements (MIPS)Software memory requirements (MB)

Flight software tradesLevel of flight software autonomyProgramming language: C or Ada

77

Flight Software Autonomy Trade

Calculated ...on EarthLow

Medium precision ….orbit propagator

Earth provides ...accurate positions

Preplanned...communications ...routing

Simple search

Calculated ...on-board ...with Earth ...input

Partial

Calculated ...on-board

Continuously ...tracks MSEs

High

GN&CCommunicationsMSE Position Determination

Level of Autonomy

Automatic ...communications ...routing

High precision ….orbit propagator

Accurate position ...calculated on-board

78

Other Flight Software Autonomy

Attitude determination and controlIncludes momentum management

Routine housekeepingThermal controlPower managementData storage

System monitoringDetects anomaliesControls safe modes

79

Flight Software Size

0

10

20

30

40

50

60

70

80

Ada C Ada C

Flig

ht S

oftw

are

(kSL

OC

)

HighPartialLow

Actual Code Code to be Developed

Some I/O device handlers can be reused

80

Flight Software Computer Requirements

0

1

2

3

4

5

6

7

8

9

Throughput (MIPS) Memory (MB)

Thro

ughp

ut (M

IPS)

/ M

emor

y (M

B)

HighPartialLow

RAD 6000 ProvidesThroughput: 10 to 20 MIPSMemory: 16 GB

Software computer requirements are met

81

Ground Software Size

Test, integration, and simulation softwareAssumed to be 4x the size of the flight softwareModeled in CERs

Initial operations softwareAssumed to be 4x the size of the flight software

0

20

40

60

80

100

120

140

160

180

200

Ada C

Gro

und

Softw

are

(kSL

OC

)

HighPartialLow

82

Software Cost

Partial autonomy with C as the programming language was chosen to meet IOC cost cap

0

10

20

30

40

50

60

70

HighPart

ial Low

HighPart

ial Low

Softw

are

Cos

t (FY

00$M

)

Initial GroundOperations SoftwareFlight Software

Ada C

83

Autonomy vs. Operations Cost

Autonomy reduces the yearly operations cost

0

10

20

30

40

50

60

High Partial Low

FY00

$M

Total Software Cost

Operations Cost perYear

84

Autonomy vs. Operations Cost

Total Software and Operations Cost for Different Autonomy Levels

25

50

75

100

125

150

175

0 1 2 3 4 5

Years after IOC

FY00

$M (n

ot in

clud

ing

infla

tion)

High Autonomy

Low Autonomy

Partial Autonomy

High autonomy is cheaper in the long run

Detailed Design:Bus Analysis

86

Bus Requirements

M008 MINERVA shall have a design lifetime of at least 6 years.

S002 Each spacecraft shall have power to support nominal operations of the spacecraft at all times, including eclipse periods.

S003 Each spacecraft mass shall not exceed 575 kg.

S011 Each spacecraft shall have the capability to boost to a disposal orbit.

B001 ADCS subsystem shall maintain pointing accuracy of 0.1 degree.

87

Bus Requirements

B002 ADCS shall provide orbit station keeping.B003 Thermal subsystem shall maintain spacecraft

components within their operating temperature ranges.

B004 Power subsystem shall provide 200 W of power during transit.

B005 Power subsystem shall provide 400 W of power throughout the operational lifetime in Mars orbit.

B006 Power subsystem shall provide 400 W-hr of energy storage.

88

Bus Requirements (cont.)

B007 Propulsion subsystem shall provide at least 2400 m/s ΔV (total).

B008 Propulsion subsystem shall provide sufficient ΔV for disposal.

B009 Spacecraft structure shall survive launch environment for a Delta III.

B010 Spacecraft structure shall survive radiation environment for the duration of the mission lifetime.

89

Bus Group Design

MATLAB software model used to perform design tradesInputs

Payload characteristicsOrbit parametersMission requirements

OutputsSpacecraft budgetsSpacecraft cost

90

ADCS Sub-System DesignDirected antenna requires 3-axis pointing stabilization

Gravity gradient/spin stabilized could not meet minimum requirements

SensorsSunHorizonGyros (safe mode)Accelerometers

ControllersReaction wheelsThrusters

91

Propulsion Sub-System Design

Launch decision allows Mars transfer ΔV to be done by launch vehicleMinimize cost - choose between EP, chemical propulsionNTO/MMH propellant

Isp = 322.5 secThrust = 4250 N

92

Thermal/Power Sub-System Design

Thermal module calculates the power needed to maintain thermal managementPower module calculates solar array area/mass based on EOL

Solar Array Flight Experiment

Batteries sized for mission life, eclipse period

Lithium-ion batteries

93

Structure Sub-System Design

15% mass margin20% structure mass factor

Power uses 30%

Payload mass calculated separately

94

Spacecraft Bus DesignSystem Component Number Mass Total Mass Total Power Critical Dim

Payload 1 37 37 190 Ant Diam = 2m

ADCS 30.7 39.0Sun sensor 6 1.2 7.0 0.8

Horizon Sensor 4 0.7 2.8 5.0Gyroscope 2 0.7 1.3 10.0

Accelerometer 2 0.1 0.2 1.2Reaction Wheel 4 3.8 15.0 22.0

Structure - 4.4 4.4 -

Propulsion 273.8 25.0Propellant - 177.4 211.8 -

Main Engine 1 4.5 4.5 15.0ACS Engine 12 0.5 6.0 -

Propellant Tank 2 10.6 21.2 - Diameter = 0.6mBlow dow n System 1 20.0 20.0 -

Feed System - 5.0 5.0 10.0Structure - 4.4 5.3 -

Thermal 7.0 11.6Heater - 2.3 2.3 11.6

Radiator - 2.3 2.3 -Insulator - 2.3 2.3 -

Power 50.1 418.0Solar Arrays 2 11.0 22.0 418.0 Area = 4.00 m^2

Electronics - 8.3 8.3 -Batteries 6 1.2 7.4 393 W-hrs

Wiring - 1.0 1.0 -Structure - 11.3 11.3 -

Launch Structure - 10.5 10.5 -

Total Mass: 409w / margin 470

Launch Vehicle Fit-Check

Four satellites fit in Delta-III fairing with 3 cm minimal clearanceBottom satellite mounts to launch vehicle adapter structureSatellite attachment rings part of satellite structure

Pyro-bolts lock rings togetherSprings separate spacecraft after rings unlock

Stowed SatelliteStowed volume ~ 4 m3

Spacecraft: Nadir Pointing SideHelix antennaHorizon sensorsPrimary sun sensorsSun-nadir steering maintains Mars-Earth-Sun pointing

Steerable main antennaSteerable solar arrays

Satellite: Internal ComponentsFirst iteration of ADCS and electronics layoutPropellant tanks shown:

NTO/MMHHe pressure regulation

Lithium/Ion batteries2 are redundantHidden in diagram

Harnessing and plumbing not modeled


Detailed Design:Operations Analysis

99

Operations Requirements

Z002 System shall have an operational lifetime of at least 5 years.

M006 At IOC the system shall be able to support at least 10 MSEs simultaneously.

S005 Constellation shall have at least 90% probability of meeting the minimum requirements throughout its operational lifetime.

S010 Each spacecraft shall have at least one recoverable safe mode.

100

Other Satellites

Operations: System Context

EarthMINERVA

Communication Positioning

Mars

• En route• On station

• MSEs• Science data

• Ground station• DSN• Launch vehicle

Requirements Requirements

• Radiation/atmosphere• MeteoritesEnvironment

Requirements

101

Operations: Functional Analysis

1. System Development

6. Conduct Training

5. Normal Operations

4. Launch and Deployment

3. Integration/ Test

2. System Production

8. Replenishment/ Replacement

7. Contingency Ops

9. Retirementor

or

102

Operations: Functional Analysis

Earth UplinkData collection/processing at EGSSegments are time/destination tagged

Mars UplinkMINERVA initiates communication per instructions

Positioning LoopMINERVA initiates positioningOn-board calculation with EGS updates

Anomaly ResolutionThree Safe Modes, Tiger Team crisis resolution

103

System Reliability: Safe ModesProgressive levels of ops reductionGraceful degradation of spacecraft and availability

Safe Mode 1: Anomaly flags or checkouts not ok, maintain high availabilitySafe Mode 2: Non-critical power or mechanical failures, EGS notificationSafe Mode 3: Critical failure, spacecraft shutdown, 14 hour self-reliance window

104

System Reliability: Failure Tree

Examination of critical failuresResult from lower level faultsMulti-path vs. complete redundancy

Setup PhaseBinary: Launch, separation, transitPartial: Detachment, deployment, capture

Normal Operations PhaseNo failureExternal: Environment, interactionsInternal: Operators, software, hardware

105

System Reliability: Event Analysis

0 1 2 3 4 5 60.75

0.8

0.85

0.9

0.95

1

M is s ion Tim eline [y ears ]

Rel

iabi

lity

[0 -

1]S y s tem Reliability over Lifetim e

4 S ats Operational A t leas t 3 S ats Operational

Time (years)

Prob

abili

ty o

f Su

cces

s

Launch toDetachment

Transit

Capture toDeployment

Normal LifetimeOperations

Detailed Design:Launch Analysis

107

Launch Requirements

Z001 System shall achieve initial operational capability by 2010.

M007 Total system mass and supporting launch structure shall be no greater than what can be launched on a single launch vehicle to a Mars transfer orbit.

S009 Launch vehicle shall be able to boost entire constellation mass to a Mars transfer orbit with a C3 energy of 6.46 km2/s2.

108

Launch Vehicle Trades

0

2000

4000

6000

8000

10000

12000

14000

Athena II Delta II Delta III Sea Launch

Payl

oad

(kg)

0

10

20

30

40

50

60

70

80

90

Cos

t ($

M)

LEO Escape Cost

109

Launch Vehicle Performance

The Delta III can provide more C3 energy than is needed for transferAdditional capability will be used to change the inclination of the parking orbit to 23.45°

Margin

Escape Performance

0

500

1000

1500

2000

2500

3000

0 10 20 30 40 50 60

C3 Energy (km^2/s^2)

Mas

s (k

g)

6.5

Detailed Design:Cost Analysis

111

Cost Requirements

Z003 At IOC the expenditures in FY2000 dollars shall be less than $300 million.

112

Cost MethodologyConcurrent engineering sessions calculated total program cost for each design iterationSpacecraft development (10% profit, 15% margin)

Design-based cost estimating relationships (CERs)*

Limitation: Accuracy of CER methodologyGround station development (10% profit, 15% margin)

Ground software x 1.5 (equipment, management, etc.)Assumption: JPL to provide space, equipment to minimize costs

LaunchDelta III launch vehicleAssumption: Reduction in Delta III costs with EELV-related efficiencies and market pressures

Transit and on-orbit operations are not included

*Applied cost factor of 1.25 (addresses uncertainty in methodology)

113

Cost and Concurrent EngineeringDesign vector

Payloadcost

SpacecraftRDT&E

SpacecraftTFU

BusPayload

Total System Development Costs

LaunchGround S/W

Systems• Margins• Factors• Learning Curve

RDT&E: Research, development, …………..test and evaluation

TFU: Theoretical first unit

114

Level of spacecraft autonomyProblem: Spacecraft autonomy drives software costsTrade space

Highly autonomous spacecraft functionsMinimal spacecraft autonomy (on-board position fix or earth position fix)

Decision: Minimal autonomy (on-board position fix)

Spacecraft propulsionProblem: Determine most cost-effective propulsion systemTrade space: Electrical versus chemical propulsionDecision: Chemical propulsion is more cost effective given launch vehicle ability to inject into Mars transfer

Major Cost Trades

115

Design Freeze Down-Select

272.228.2244.0Delta IIChem3*Option 4

290.929.8261.2Delta IIIEP3*Option 3

297.931.5266.4Delta IIIChem4Option 2

312.333.3279.1Delta IIIEP4Option 1

Total($M)

Margin*(.15)

Cost($M)

LaunchVehicle

PropSystem

# of S/C

* Does not meet all performance requirements (coverage, Gb/sol)* On spacecraft and ground station development costs. No margin on ...launch costs.

116

Spacecraft Cost Model

CERs from SMADAssumes deep space and Earth orbiting systemsAccuracy to within 25-50%

Calculate RDT&E, TFU cost separatelyTFU cost scales with number of spacecraft according to learning curve

117

Major Elements of Cost

Margin (11%)

Launch (19%)

Spacecraft (59%)

Ground Station (12%)$31.5 M

$56.3 M

$39.8 M

$170.3 M

118

Life Cycle Costs

Spacecraft

Launch Ground

Margin

Operations (5 years)• $129.0 M

Operations (transit)• $20.2 M

Total Life Cycle Cost (5 year mission): $447.1 M

System Level Issues

120

System-Level Risk Management Strategy

Cost Risk (Medium)Source: CER methodology; software & launch costsStrategy: Apply cost factor (1.25) and hold margin (15%)

Technical Risk (Low – Medium)Source: Mission integration, software development, cross-linksStrategy: Maximize use of proven hardware and software

Schedule Risk (Low)Source: Complexity of deep space programStrategy: Hold margin before 2007 launch window

Maintain low risk through cost and schedulemanagement and reliance on existing technology

121

0807060504030201

IOC

On-orbit checkout

Mars Transfer

Launch

Launch Site Ops

Integration & Test

Flight Software

Fabrication

Design

CY

MarginMargin

Program Schedule

122

Total Program Cost: $297.9 M*Total Program Cost: $297.9 M*

Program Yr

4.84.826.126.151.151.167.067.067.067.051.151.126.126.14.84.8Funding

20082007200620052004200320022001Calendar Yr

87654321

Profile Cumulative

(CY00 $M)

*Includes 15% margin (Note: CER methodology limits validity of cost estimate)

$0$20$40$60$80

1 2 3 4 5 6 7 8

Program Year

$0$100$200$300

1 2 3 4 5 6 7 8

Program Year

Funding Profile

123

MINERVA Science Capabilities

Improve Mars gravity field modelIndirect gravitational study of Phobos and Deimos

Atmospheric composition of MarsAbsorption and scattering properties of Martian atmosphere

Radio scienceStudy solar corona and interplanetary medium

124

Post-IOC System Expandability

Upload software with improved autonomyProvide positioning and communication service to other spacecraftRelay between MSEs without Earth interactionAutomate ground operationsAdd more spacecraft to constellation

Improve coverage, availability, and reliabilityInclude upgraded capabilities (e.g. remote sensing)

Replenish constellation as spacecraft fail

125

Lessons LearnedMethods for discovery of errors and disconnects

Usefulness of frequent integration meetings and status briefingsEvaluation of concurrent engineering session results

TransitionsTeam structure changed after TARR, delaying some tasksPost-PDR transition much more rapid, effective

Concurrent engineeringUseful for rapid characterization of design options via real-time inter-team communicationMust be supplemented with detailed design analysis between sessionsICEMaker is useful interfacing toolMore automation would speed process

Backup Slides

Backup Slides:Orbit Analysis

128

Transit Overview

Departure burn 18 Aug 07, 09:56Separation 18 Aug 07, 13:25Deploy arrays 18 Aug 07, 13:31Initial checkout 18 Aug 07, 14:00Exit Earth SOI 21 Aug 07, 02:35Arrive Mars SOI 29 May 08, 10:56Circularization 03 Jun 08, 18:18Deploy antenna 03 Jun 08, 18:20Test/calibration 09 Jun 08, 22:20IOC 10 Jul 08, 00:00 Earth (start) Mars (start)

Earth (final)

Mars (final)

Separation &Deployment

Spin-up &Insertion

CorrectionBurn

AlignmentBurn

129

Percentage of Time in ViewSingle Satellite Failure

Constellation provides >50% coverage in the ± 150 latitude bandReduced coverage

up to ± 650

130

Revisit TimeSingle Satellite Failure

The maximum time between satellite passes is <100 minThe average time is <45 min

131

Contact DurationSingle Satellite Failure

On average, a satellite will remain in view for 50 minutes.

Backup Slides:Payload Analysis

133

Link MarginsEarth - MINERVA link:

Uplink: 28.8 dB, downlink: 3.09 dBMINERVA - Mars link:

Uplink: 5.29 dB, downlink: 4.73 dBMINERVA cross-link:

Uplink and Downlink: 17.4 dBMINERVA cross-link with Ka-band for DTE link:

Uplink: 16.65 dB, downlink: 2.97 dBMINERVA cross-link with omni-directional antenna for case of the loss of attitude control:

Uplink and Downlink: 12.4 dB

134

Communications Analysis:Worst Case

Two MSEs on the dark side of Mars.Each of the MSEs is at the edge of the cone of MINERVA-Mars link.Each MSE has no more than 10W RF power.Largest distance between Earth and Mars is equal to 401,300,000 km.Maximum distance between MINERVA satellites is equal to 7,633 km.

135

Payload Electronics Hardware

3 amplifiers (total output power ≈165 W)2 Ka-band and X-band supporting transponders2 computers1 UHF transceiverOne ultra-stable oscillator

One failure of a critical component(amplifier, transponder, computer)

≠loss of the satellite

136

Frequency Used For Future Mars Missions (from Chad Edwards speech)

137

High gain antenna failure• One antenna failure:

• Still fully meet the requirements

• More antenna failures:

• Graceful degradation of performance

Failure Mode Analysis

.

DSN

138

Cross-link antenna failure• If one antenna on a satellite fails:


• If more antennas fail:



.

DSN

139

UHF antenna failure• One antenna failure:


• More antenna failures:



.

DSN

140

Accuracy Over Time

141

Positioning Performance

First estimate accuracy depends on geometry w.r.t. satellite ground trackTime to reach accuracy is a function of

Orbital inclination MSE latitude

Best performance around the equator (coverage)

142

Positioning Performance

Comparison with 30 degrees inclination:

EL/KM

143

Time to Get 100 m Accuracy: Comparison with 30° inclination

Probability to reach 100 m accuracy (1 σ) within certain time:

0° latitude

144



15° latitude

145


[min]


0° latitude

146


Probability to reach 100 m accuracy (1 σ RSS) within certain time:

15° latitude

147

Software Cost

Ada

Ground SoftwareFlight SoftwareCost per SLOC

$ 435

C

$ 220

$ 220$ 726

Software cost estimated by SLOC

148

Computer Hardware - RAD 6000Radiation hardened version of IBM Risc 6000 Single Chip CPU (32 bit)

Chip dimensions: 8” x 9” x 2” inches

Mass: ~5 kg

Memory: 128 MB of DRAM + 16 GB of EEPROM

MIL-STD-1553 interface

Processing speeds20 MHz (22 MIPS) using 9 W10 MHz it (11 MIPS) using 5.5 W 2.5 MHz (2.7 MIPS) it uses 2.5 watts.

Two processors (2 for 1 redundancy)

Backup Slides:Bus Analysis

External Satellite Components

Internal Satellite Components

Backup Slides:Launch Analysis

153

Launch Vehicle PerformanceLEO Performance

0

2000

4000

6000

8000

10000

0 3000 6000 9000 12000 15000

Altitude (km)

Mas

s (k

g)

Escape Performance

0

500

1000

1500

2000

2500

3000

0 10 20 30 40 50 60

C3 Energy (km^2/s^2)

Mas

s (k

g)

Backup Slides:Operations Analysis

155

Functional Flow1 System

Development2 System

Production3 Integration/

Test4 Launch andDeployment

5 Normal Ops

6 ConductTraining

7 ContingencyOps

8Replenishment/

Replacement9 Retirement

1.1 ConceptDevelopment

1.2 DetailedDevelopment

1.3 OpsDevelopment

2.1 Build

2.2 Gather OpsStaff

3.1 Hardware

3.2 Software

3.3 Ops

4.1 LaunchScheme

4.2 Space Flight 4.3 TransMarsInsertion 4.4 Deployment 4.5 Checkout 4.6 Checkout

Transmission

4.5.1Subsystems 4.5.2 System

4.5.2.1Determine

Initial Position

4.5.2.2Manuever IfNecessary

9.1 Notice toUsers

9.2 Shutdown 9.3 Disposal

7.1 PerformSystem Checks

7.2 AnalyzeChecks

7.3 Go toAppropriateMode (Safe

Mode)

7.4 FixAnomoly

7.5Communicate

Anomoly

7.1.1 KnowNominal State

7.1.2 DetectAnomolous

State

5.1 Nav 5.2 Comm 5.3 (Obs) 5.4 System

5.1.1 CalculatePosition

5.1.2 ProvideNav Solution

5.1.1.1 ReceiveEphemeris

5.1.1.2 UpdatePosition From

Previous

5.1.1.3Manuever ifNecessary

5.1.2.1 SendTwo-way

Range to MarsUnit

5.1.2.2 ReceiveRange Signal

Back

5.1.2.3Calculate Mars

unit Position

5.2.1 AcquireTransmission

5.2.2Determine

Comm Scheme5.2.3 Retransmit

5.2.1.1 EGSSignal to DSN

5.2.1.2 DSNSignal toMinerva

5.2.1.3 MarsUnit Signal to

Minerva

5.2.3.1 SendWithin

Minerva5.2.3.2 Send to

Other MarsUnits

5.2.3.3 Send toDSN

5.2.3.4 Send toEGS

5.4.1 ProvideInfrastructure

to SupportPayload

5.4.1.1 SystemReliability

5.4.2 CostEffective

or

or

or

or

or

and

andandand

or

or

156

Earth Uplinkcollect data/commands

from PI for Mars Units atEGS

collect data/commandsfrom PI for MINERVA at

EGS

generate EGSdata/commands/updates/ephemeris

at EGS

transmit to DSN(assumed access)

DSN transmit toMINERVA

MINERVA checkstransmission

MINERVAde-interleaves signal

segments forretransmit sent to

buffer

segments forMINERVA sent to

computers

segment is storeduntil time tag directs

acquire contact withMars Unit

acquire contact withMINERVA crosslink

satellite

transmit to Mars Unit

.

...

EGS data processing:interleaving, time tagging,

destination

signal terminates atMINERVA crosslink

satellite

receive confirmationfrom Mars Unit

receive confirmationfrom crosslink satellite

MINERVA associates listof users (comm and

positioning)

MINERVA updatesposition from

ephemerismaneuver if necessary

or

or

and

or

157

Mars Uplink

MINERVA sends comminitialization signal to

userMINERVAclears buffer

MINERVAreceives user signal data stored in buffer MINERVA sends

confirmationMINERVA interleaves datawith next transmission to

DSN

MINERVA receivesconfirmation from EGS

(through DSN)

158

Positioning Loop

MINERVA updatesposition from on-orbitpropagation analysis

MINERVA sendsinitialization signal to

userMINERVA

receives user reply

MINERVA calculatespositioning solution

MINERVA sendssolution (for an allotted

time)MINERVA endspositioning loop

159

Anomaly Resolution

MINERVAsubsystem checkout

not OK

MINERVA subsystemsends anomaly flag

go to Safe Mode 1 run autonomousanalysis

go to Safe Mode 2if necessary

go to Safe Mode 3if necessary

send Safe Modenotification to EGS

receive EGS SafeMode response

implement EGSinstructionsor

fix anomaly(correct, reroute)

or

or

160

Failure Tree: Setup

launch failure

successful launch

separation failure

successful separation

detachment failure

successful detachment

1 to 3 successfullydetach

transit failure

successful transit

capture failure

1 to 3 capturesuccessfully

successful capture

deployment failure.

1 to 3 successfullydeploy.

successful deployment.- pyros - mechanics- power - propulsion- to correct altitude

- thrusters- correct altitude- enter correct orbit- enter correct spacing

- solar arrays - antennas

MINERVA Setup

- computers

deployment failure

1 to 3 successfullydeploy

successful deployment

161

Failure Tree: Normal Lifetime Ops

Lifetime Ops

no failure

externally-causedfailure

internally-caused failure

radiation

meteorites

operators

software

hardware

improper command

fault/data oversight

improper code

inability to compensatefor input/unknown

battery failure

engine failure

wiring failure

main computer failure

data hard storage failure

data soft storage failure

thermal cooling failure

propellantcontainment failure

attitude sensor failure

control actuator failure

antenna failure

power supply failure

transponder failure

162

Reliability (and Failure Rates)Launch: 0.997 (or 0.90)Separation: 0.99Detachment: 0.99Transit: (0.005 failures/year)Capture: 0.99Deployment: 0.99ADCS: (0.001 failures/year)Payload: (0.00201 failures/year)Power: (*** failures/year)Propulsion: (0.005 failures/year)Thermal: (0.002 failures/year)Computers: (0.005 failures/year)

Backup Slides:Cost Analysis

164

Problem: Spacecraft autonomy drives software costsTrade space:

Highly autonomous s/c functionsFlight software: $24.8MGround software: $50M

Minimal s/c autonomy (on-board position fix)Flight software: $17.6MGround software: $20.5M

Minimal s/c autonomy (Earth position fix)Flight software: $16.4MGround software: $19.1M

Decision: Select minimal autonomy (Earth position fix) due to program cost constraints

Cost Trade: Level of Autonomy

165

Notes on Concurrent Engineering

Design sessions enabled thorough exploration of trade space via real-time inter-team communication

Earth parking orbitConstellation altitude# s/cOrbit inclination

ICEMaker is useful interfacing toolMore automation would speed process

Models in ExcelMatlab/Excel integration

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